This invention pertains to airfoils in gas turbines and in particular to cooling of the airfoil trailing edge.
Airfoils of gas turbines, turbine rotor blades and stator vanes, require extensive cooling in order to prevent damage due to overheating. Typically such airfoils are designed with a plurality of passages and cavities arranged spanwise from the root to the tip of the airfoil for cooling fluid to flow through. The cooling fluid is typically air bled from the compressor having a higher pressure and lower temperature compared to the gas traveling through the turbine. The higher pressure forces the air through the cavities and passages as it transports the heat away from the airfoil walls. The cooling air finally leaves the airfoil by means of exit slots along the sidewalls and at the trailing edge of the airfoil. The cooling of the airfoil material by a cooling construction of passages and cavities occurs by several different physical means such as film cooling, impingement cooling, and diffusion.
The trailing edge of the airfoil is particularly difficult to cool for several reasons. First, the cooling air has an increased temperature by the time it reaches the trailing edge reducing the temperature difference between it and the material to be cooled. Second, the material of the trailing edge is relatively thin such that it is especially susceptible to thermal stress and damage due to overheating.
Furthermore, efforts have been made to minimize the thickness of the trailing region in order to improve the aerodynamics of the airfoil. A thin trailing edge presents less blockage to the working fluid and as such reduces the pressure loss allowing a greater turbine performance. Hence, as the airfoil""s aerodynamics is improved by the reduction of the edge thickness the cooling of the airfoil material becomes a greater challenge.
The German patent DE 25 55 049 (U.S. Pat. No. 4,153,386) discloses an example of an airfoil with a trailing edge having a passageway for cooling fluid leading from an internal cavity to an exit slot. Rectangular corners shape the end portion of the trailing edge with the exit slot. The finite thickness and sharp corners of the end portion cause vortices about the exit slot so that the cooling effectiveness is compromised. Subsequently, a large amount of cooling fluid is necessary in order to effect a sufficient reduction of the edge temperature.
U.S. Pat. No. 4,601,638 discloses an improved design for the trailing edge of a gas turbine airfoil for an aero engine with spanwise arranged channels leading from internal cavities to exit slots. The exit slots for the cooling fluid are arranged on the pressure side of the airfoil. Here a so-called cutback distance is defined as the distance from the tip of the airfoil to the edge of an exit slot on the pressure side that is farthest away from the airfoil trailing edge. The design includes in particular a ratio t/s of the pressure side lip thickness t to the width of the cooling fluid channel s that is equal to or less than 0.7. For a blade with a given nominal pressure side lip thickness t of 0.035 inches and a tailing edge thickness d of 0.08 inch this ratio t/s results in a cutback distance equal to or greater than 0.24 inches. In other words, the constraint on the ratio t/s results in a constraint on the minimal cutback distance.
U.S. Pat. No. 6,004,100 discloses another cooling construction for a gas turbine airfoil trailing edge in an aero engine with similar cavities and channels leading to cooling ports on the pressure wall side through which the cooling fluid exits the airfoil. The channels leading from an internal cavity to the cooling port each have a specially designed cross-section comprising a pair of fillets. In a further development the channels each comprise a jog in order to prevent high cycle fatigue along the narrow edge. The design allows in particular the cutback distance to extend all the way to the tip of the airfoil. The thin resulting walls can lead to problems in production of the airfoil.
It is the object of this invention to provide an airfoil for a gas turbine with a cooling construction for its trailing edge by which component life as well as turbine performance are improved over the state of the art.
In particular, it is the object of this invention to provide an airfoil with a minimized edge thickness and with a cooling construction for its trailing edge that sufficiently cools the trailing edge with a minimized amount of cooling fluid.
An airfoil attached to a rotor in a gas turbine comprises a pressure sidewall and a suction sidewall each with an inner and outer surface. One or more cavities are arranged spanwise, and/or radially, between the two suction and pressure sidewall. A cooling fluid bled from a cooling fluid source flows through the cavities. Furthermore, a plurality of exit passages is arranged within the airfoil, which lead from the cavities to exit ports positioned on the pressure sidewall and near the trailing edge of the airfoil. The exit passages are delimited in part by the inner surfaces of the suction side wall and a pressure side lip, that is the end portion of the pressure side wall that leads up to the exit ports at the trailing edge. The suction sidewall extends such that it forms the tip of the trailing edge of the airfoil.
According to the invention the suction sidewall is shaped such that the tip of the airfoil trailing edge has a cross-section that comprises a part of a circle whereas the part of the circle is equal to or less than a semi-circle. The part of the circle extends from the suction side of the airfoil around the tip of the airfoil to the pressure side of the airfoil and beyond into the exit passage. The exit passages each comprise a bend in the region of the airfoil trailing edge tip toward the pressure side where the exit passages form an exit plenum for the cooling fluid to diffuse and cool the airfoil trailing edge.
The invention is thus based on a multi-diffusion concept, and is produced by the diffusion of the cooling fluid flow form the several exit passages into the large exit plenum about the bend region of the several exit passages. This is in particular realized by the rounded shape of the tip region according to the invention of the suction sidewall of the airfoil.
The tip of the airfoil trailing edge according to the invention has a cross-sectional shape in the likeness of a nose. Cooling fluid flows from the internal cavities through the exit passages towards the exit ports. As it reaches the region of the nose-shaped tip it also encounters the bend of the passages, which is delimited by the inner surface of the suction sidewall. The cooling fluid impinges on the inner surface of the suction sidewall and cools by impingement cooling the trailing edge tip region of the airfoil. Following the bend of the inner surface of the suction sidewall it flows toward the exit port on the pressure sidewall. It flows into an exit plenum in the region of the bend and the pressure side lip, diffuses and cools by film cooling the rounded the tip of the airfoil. The tip of the trailing edge is thus cooled by impingement as well as by film cooling.
In the longitudinal direction of the airfoil a multitude of exit passages lead onto the exit plenum on the pressure side of the airfoil. The exit plenum on the pressure side is enlarged compared to cooling constructions of the state of the art and thus allows a greater diffusion of the cooling fluid. Due to the relatively large size of the exit plenum the velocity of the cooling fluid flow is reduced which enhances the film cooling effectiveness. The nose-like shape of the airfoil trailing edge tip further effects that the cooling fluid follows the surface of the airfoil tip while cooling it whereas few or no vortices can form which would impede the cooling effectiveness. This further improves the film cooling effectiveness and as a result lowers the amount of necessary cooling fluid.
The cooling of the airfoil trailing edge is enhanced by film cooling of the tip on one hand and by impingement of the inner surface of the tip on the other. This enables the construction of the airfoil trailing edge with a smaller thickness by which the aerodynamics of the airfoil and as a result the performance of the turbine are improved.
Furthermore, there is a cutback distance equal to the distance between the tip of the trailing edge and the end of the pressure side lip. This design of the exit passages with a bend and a rounded shape of the tip effectively shortens the cutback distance. Calculations have shown that a reduction of the cutback distance improves the film cooling effectiveness. As a result, this feature reduces the amount of necessary cooling fluid.
In a particular embodiment of the invention the exit passages leading to the exit ports have a cross-section perpendicular to the cooling fluid flow that is circular, rectangular with rounded corners, or oblong. The sidewalls of the passage are diffused in the direction of the cooling fluid flow such that the flow velocity of the cooling fluid is further reduced and the film cooling effectiveness is enhanced.
In a particular embodiment of the invention the diffusion angle of the exit passage is less than three degrees. Calculations have shown that for this amount of diffusion the cooling effectiveness is optimized. In particular, for the range of the ratio of the length to the diameter of the cooling channel that is given for a typical airfoil a diffusion angle of less than three degrees prevents internal flow separation and a higher loss coefficient.
The exit passages lead onto a large exit plenum. For structural strength of the airfoil supporting ribs are arranged in this region substantially parallel to the exit passages. The supporting ribs are arranged every two or more exit passages. For example, four exit passages extend to the bend and onto pressure side of the airfoil whereas the first and fourth exit passages are adjoined on one side by a supporting rib that extends to the tip portion of the trailing edge of the airfoil. The large exit plenums are formed between such supporting ribs. In these exit plenums the cooling fluid can diffuse not only in the direction toward the trailing edge tip but also in the radial directions. Within this exit plenum a mixing of the cooling fluid with mainstream gas flow and the formation of vortices are mostly prevented enhancing the film cooling effectiveness.
In a particular embodiment of the invention the size of the cutback is less than 0.12 inches. This shortened cutback distance contributes to the increased film cooling effectiveness. Furthermore, the current invention yields a smaller cutback with a larger trailing edge thickness.
In a further embodiment of the invention the airfoil surface is provided with a protective coating. The coating is applied to the outer surface of the pressure sidewall and the suction sidewall. The coating serves as protection against erosion, corrosion and oxidation. It is applied with a large thickness of 0.014 to 0.018 inches, which provides a longer oxidation life and hence a significantly greater component life as the component life increases exponentially with the coating thickness.
The thickness of the airfoil trailing edge is reduced compared to designs of the state of the art. A thin trailing edge thickness according to the invention can accommodate a thick protective coating while the thickness of the resulting trailing edge remains still smaller than those of the state of the art and presents reduced blockage to the gas flow.
Furthermore, the design according to the current invention has a relatively large exit plenum such that the protective coating can be accommodated without blocking and impeding the flow of the cooling fluid out of the exit port.
In a particular embodiment of the invention the pressure side lip thickness is equal to or greater than 0.035 inch and the trailing edge diameter, that is the diameter of the nose-shaped airfoil tip, is equal to or greater than 0.08 inch.
In a further particular embodiment of the invention the exit port of the exit passage has a width equal to or less than 0.08 inch.
In a further particular embodiment of the invention the ratio of the pressure side lip thickness to the height of the cooling fluid exit passage is greater than 0.7.